The present embodiment of the invention relates generally to transonic structures and more particularly to a device for preventing abrupt loss of lift where a boundary layer tends to interact with shocks during transonic flow conditions.
Predictable flight performance characteristics are of paramount importance in aircraft design. In flight, if lift suddenly decreases on one wing relative to the other, then a loss of control results and corrective action by the pilot is required. At low altitudes, if the wing abruptly losses lift, and corrective action cannot be quickly implemented, then a loss in aircraft and pilot may result. In military aircraft, predictable flight performance characteristics are necessary to precision maneuvers such as formation flying, and combat. In particular, military aircraft performance requires predictable flight performance characteristics during high-speed flight.
During high-speed transonic performance testing of an aircraft, many maneuvers are flown to test the capability of the aircraft. During such maneuvers a phenomenon known as Abrupt Wing Stall may occur. An abrupt asymmetric wing stall of the aircraft wing causes wing-drop. The abrupt loss of lift on a wing causes the aircraft to lose lateral control and the aircraft to roll uncontrollably about its longitudinal flight axis. Total loss of lift on both wings results in an aircraft stall that is manifested by the aircraft rotating forward and falling. In addition, flight control surfaces such as ailerons and flaps, which are located in the regions of flow separation, lose effectiveness.
For the past half century, abrupt wing stall phenomenon has been known to occur in many military combat aircraft such as the F/A-18E/F. The F/A-18E/F is a military aircraft that embodies enhancements to the original F/A-18 fighter-attack aircraft. The wing of the F/A-18E has a non-constant chord leading edge flap design. The wing includes an outboard leading edge flap having a snag to be described hereinbelow. During development of the Navy""s F/A-18E, a complex shock-boundary layer interaction developed during transonic flight. A localized leading edge flap separation occurred just inboard of the snag and interacted with a preexisting shock on the upper wing surface, thereby causing wing-drop. Wing-drop was most likely to occur when the airplane angle of attack was increased as typically occurred during in-flight maneuvers at transonic velocities.
As is well known in the art, flow separation tends to occur at projections or structures located on the wing such as at a joint between a main wing and an engine nacelle. Such a phenomenon is referred to as flow interference. Use of fillets between the wing and the nacelle has been successful to prevent flow separation. However, fillets result in increased frontal area of the nacelle and wing combination thereby increasing form drag especially at higher speeds, in particular in and above the transonic range. xe2x80x9cForm dragxe2x80x9d is friction caused by the flow of air molecules past the skin of the airplane including all of the aircraft structures in contact with the airstream. Form drag produces a force that opposes the velocity of the airplane, thereby requiring more thrust from the engines. The application of these fillets is not generally suitable for stabilizing recovery shocks.
Vortex generators mounted on the upper surface of a wing have also been used successfully to prevent flow separation. Although vortex generators can be easily installed on existing structures, vortex generators are comparatively ineffective and incapable of preventing flow separation in regions of very violent flow as would be experienced in a highly maneuverable fighter aircraft especially aircraft operating at transonic speeds. Furthermore, the application of vortex generators is not generally applicable to stabilizing flow recovery shocks.
The use of a porous surface positioned at the upper surface of the wing has been used successfully to fix the position of the recovery shock. By fixing the shock, the region of shock induced flow separation remains fixed on the wing thereby preventing movement of the recovery shock from affecting lift patterns on the wing. However, porous surfaces have a detectable radar signature, which is not desirable in military aircraft. In addition, porous surfaces cause the recovery shock to be fixed forward at lower speeds resulting in lower coefficient of lift (CL). In addition, porous surfaces cause earlier flow separation, which leads to early buffet onset, dramatically fatiguing the pilot and aircraft. Furthermore, the buffeting tends to induce vibration on the wings and structure of the aircraft leading to premature structural failure. As is well known in the art, Lift is related to the lift coefficient CL by the relationship L=xc2xdxcfx81v2 CL, where CL is generally linearly proportional to angle of attack xcex1. xe2x80x9cxcex1xe2x80x9d is the angle between the free stream velocity vector V∞. and the wing chord or other relevant point of reference on the aircraft. The xe2x80x9cwing chordxe2x80x9d is defined to be a straight line drawn from the leading edge of the wing to the centroid of the trailing edge of the wing. Therefore, when porous surfaces are used, higher angles of attack a are required to fly the airplane in order to compensate for the reduced CL resulting in reduced visibility and increased aerodynamic drag. Aerodynamic drag is also proportional to the angle of attack xcex1. Drag is related to the drag coefficient CD by the relationship D=xc2xdxcfx81v2 CD , where CD is generally linearly proportional to angle of attack xcex1.
Accordingly, there is a need to reduce or eliminate abrupt wing stall phenomenon.
There is a further need to have a minimum radar detection signature.
There is additionally a need to minimize vibration due to fluid dynamic conditions.
There is yet another need to minimize the reduction in Lift and the increase in Drag.
The present invention is directed to a device that satisfies these needs. A stabilizer having features of the present invention comprises an inboard end, and an outboard end that forms a local airfoil. The outboard end is positioned opposite the inboard end. An upper surface extends between the inboard end and the outboard end. A lower surface also extends between the inboard end and the outboard end, and is positioned to oppose the upper surface. A leading edge between the upper surface and lower surface forms the stabilizer nose. A trailing edge forms a generally concave surface and is positioned opposite the leading edge. The outboard end of the stabilizer forms a predetermined angle omega with the trailing edge for positioning the outboard end relative to the trailing edge. The outboard end of the stabilizer forms a predetermined angle tau with the leading edge for providing forward sweep angle.
A vehicle having the features of the present invention has a main body and at least one reaction body. The reaction body comprises a reaction body leading edge and, a snag extending from the reaction body leading edge. The reaction body is attached to the main body to produce a reaction force. At least one stabilizer for stabilizing a shock formed at transonic speeds has an inboard end and an outboard end. The outboard end forms a local airfoil and is positioned opposite the inboard end. The stabilizer has an upper surface extending between the inboard end and the outboard end, a lower surface extending between the inboard end and the outboard end and, the lower surface opposing the upper surface. A stabilizer leading edge forms a stabilizer nose disposed between the upper surface and the lower surface. A trailing edge opposes the leading edge. The outboard end of the stabilizer forms a predetermined angle omega with the trailing edge of the stabilizer for positioning the outboard end of the stabilizer. The outboard end of the stabilizer forms a predetermined angle tau with the leading edge for providing forward sweep angle. The trailing edge of each stabilizer is attached to the reaction body leading edge inboard of the snag.
The vehicle encompasses any vehicle that travels through a fluid such as air, water, plasma or other fluid. The stabilizer may be positioned upon any reaction surface including a wing, fin, sailplane, leading edge flap, rudder, plane or other control surface. Where the main body has characteristics of a reaction body then the stabilizer may be attached to the main body. Installation of the stabilizer upon military aircraft is within the contemplation of the invention. When installed upon the leading edge of a reaction body, one embodiment of the stabilizer contemplates a forward sweep angle equal in magnitude to the sweep back angle of the opposing wing in order to minimize radar signature.
Accordingly, it is an advantage of the present invention that abrupt wing stall phenomenon is reduced or eliminated. It is a further advantage that radar detection signature is minimized. It is yet another advantage that vibration due to fluid dynamic conditions is minimized. There is yet another advantage in that the reduction in Lift and the increase in Drag are minimized.